The term "fly-by-wire" has become recognized in the aircraft control art as describing systems for aircraft control using electronic or electrical control paths. That is, electronic or electrical control paths replace the mechanical or hydraulic control paths of previous aircraft control systems. Although it has been recognized for sometime that electrical control systems offer advantages over previous mechanical or hydraulic controls, there has been a reluctance to utilize electronic systems because of the belief that a mechanical linkage, for example, provides a more reliable system. Thus, most fly-by-wire control systems employ multiple signal paths each independently capable of carrying control signals for safe operation of the aircraft. If one or more of the electrical control paths is damaged, the remaining paths will function to provide control signals for operation of the aircraft. This need for redundant control is particularly important in military aircraft where battle damage may disrupt one or more of the control paths. For additional protection against battle damage, it is often a requirement of aircraft specifications that each of the multiple control paths be run through different parts of the aircraft.
Fly-by-wire control systems have received much recent attention in research and development efforts throughout the industry. The desire for improved survivability is a strong incentive for this effort. Also, more precise control for nap-of-the-earth flight and the desire to couple guidance and navigation controllers to the actuation system of an aircraft makes fly-by-wire techniques assume additional advantage over mechanical or hydraulic control systems.
Typical of fly-by-wire control systems is that described in U.S. patent application Ser. No. 971,712, filed Dec. 21, 1978, for helicopter control. While the present invention is directed primarily to helicopter control, it will be understood that the invention is applicable to all aircraft control.
It was early recognized that fly-by-wire control systems must be fail operate, that is, the system must continue to operate after multiple path failures. In the systems hereinafter described, the actuation system includes four control loops and the system must continue to operate after two control loop failures. This is to provide reliability such that the vehicle (aircraft) be substantially immune to catastrophic failures which might result in loss of life and/or destruction of the aircraft. In the system to be described, to provide "two-fail-operate" capability the system includes redundancy of electrical and hydraulic power supplies, sensors, electronics, signal paths, and actuating devices.